It is well known in the aircraft engine field that improved thrust-to-weight ratios are a desirable achievement in the fabrication of jet engines. To improve the ratio either the thrust can be increased or the weight of the engines can be reduced. To reduce the weight of the engine either lighter alloys having the same strength must be substituted for presently used alloys or alloys which are basically stronger than those presently employed must be substituted for the presently used alloys.
Improved thrust can also be achieved by operating the engine at higher temperatures. It is known that in general jet engines operate at higher efficiency if they operate at higher temperatures. If the operating temperature of a jet engine can be increased by 100.degree. the efficiency of operation of the engine can be significantly improved. Jet engines last more than 10 years in service. If the fuel consumed by a jet engine is reduced by a significant degree over the 10 or more years of expected life of a jet engine, then there is a cost saving in the operation of the engine which is very substantial and which permits the engine to be formed at higher costs. The higher engine cost is more than offset by the lower costs of operation of the engine.
Generally speaking alloys must be capable of operating at relatively high temperatures to be employed in the hot portion of a jet engine. Up to now the highest temperature of metal parts in a jet engine is about 2200.degree. F. The present invention is directed toward alloy compositions and articles capable of operating at temperature up to 2200.degree. F. but which have substantially reduced weight relative to those presently used in engines.
In determining the value of a weight reduction in improving the thrust-to-weight ratio of a jet engine, consideration must be given as to where the alloy will be employed in a jet engine. If it is employed in the static portions of the engines then the difference in weight of the engine parts which is achieved by substituting the lower weight parts for the heavier weight parts gives a direct indication of the weight reduction.
By low stress as used herein is meant physical force applied to the articles. Thus while the article may be employed at high temperatures and may have a high thermal gradient, extending from one portion of the article to another, there is no high level of physical force applied to or developed in the article as, for example, a high centrifugal force due to rapid rotation of the article.
More particularly the present invention relates to methods for forming parts for aircraft engines for use at high temperatures. It relates, for example, to nozzle guide vanes which are subject to use at high temperatures but which are not subject to high stress as from impact of other articles at the high temperatures. It relates as well to individual articles which are formed.
It is known that many aircraft engine components such as flaps, seals, vanes, and the like must endure very high temperature as the are exposed to very high temperature gases of combustion within the jet engine. Such parts endure some pressure as they are located in regions of the engine where the gases which are undergoing combustion deliver very high levels of heat to surrounding surfaces. Some of the parts, including flaps, seals and vanes, are employed to direct the movement of the gases undergoing combustion within the jet engine.
In the first stage turbine of an aircraft engine, the vanes are the stationary components of a nozzle which guides the hot gases from the combustor against the rotating turbine blades. The rotating turbine blades extract the energy needed to power the aircraft. Since the vanes are immediately behind the combustor, they are exposed to extremely high gas temperatures. Because these gas temperatures are greater than the melting temperature of the vane alloys, the vane must be cooled by lower temperature gases. The pressure differential between the combustion gases on the exterior of the vane and the cooling gases on the interior of the vane will load or stress the wall of the vane. This loading is small but may be great enough to cause creep, or eventually, rupture, of the vane wall. Failure of the vane also may occur due to thermal fatigue cracking. The thin leading and trailing edges of the vane tend to be heated and cooled faster than the central region of the vane as the engine temperature is changed throughout the operating cycle. The central region of the vane will act to constrain the edges, leading eventually to plastic deformation and thermal cracking. Cracking may degrade the cooling efficiency and may allow temperatures locally to rise, leading to accelerated oxidation or, in the extreme, to melting.
Exhaust nozzle components, such as flaps and seals, also are subjected to low mechanical forces, but can experience large thermal strains in the engine environment. These components are located in the after-burner regions of high-performance aircraft engines. When extra thrust is required during operation, fuel is added to the engine exhaust to promote additional burning. Exhaust nozzle components typically fail via a hot-streak-induced thermal fatigue mechanism in the superalloy face sheet. The hot streaks in the exhaust gases cause large thermal gradients in the face sheet; the resulting thermal stresses cause yielding and permanent deformation of the face sheet. The thermally distorted face sheet seals poorly, causing cold air leaks from the higher pressure air outside the nozzle, further intensifying the thermal gradients. The cyclic thermal gradients cause thermal fatigue cracking and ultimately component failure.
While the above components are not subjected to high stress they are, as indicated, subject to high temperature and they also may be nested in the engine structure so that one part must move relative to other parts without interference. Accordingly any distortion of the parts either due to the method of fabrication or in the use and operation of the components is detrimental to the efficient control of movement of the jet of gases through the engine and also of the efficiency of the operation of the engine.
Another problem encountered in the use of such engine components is failure primarily due to thermal fatigue cracking. The cracking can produce the same results as the distortion of the parts and prevent their easy and efficient nesting and movement for control of the dimensions of the gas jet within the engine and the efficient flow of gases through the engine.
It is known that the refractory metals and notably chromium, molybdenum, tungsten, niobium and tantalum have very high melting temperatures and also have low coefficients of thermal expansion. However, these same materials (except for Cr) have just about no resistance to environmental attack at temperatures above 2000.degree. F. Some coatings have been developed for these materials but the components made with the coated metals have had problems due to poor thermal expansion matching of the coating versus the substrate metal and accordingly have been marginally suitable for the severe thermal cycling exposure which is inherent in the components of aircraft jet engines.
However, in dealing with the rotating parts of the engine the significance of weight reduction can be increased particularly where the weight reduction is achieved at the outer rim or outer regions of the rotating parts. One reason for this result is that the rotating parts of a jet engine rotate at very high velocity. In present engines the parts actually rotate at about 12,000 revolutions per minute. When a heavier part is held at the outer edge of a disk the disk itself must be stronger in order to hold the heavier part during its rotation at about 12000 revolutions per minute. A relatively small reduction in the weight of that part can lead in turn to a reduction in the size and strength of the disk which holds the part. This is because the part itself when rotated at 12,000 revolutions per minute develops a much higher centrifugal force than a lighter part. The higher centrifugal force must be opposed by a higher tensile force through the disk. It has been determined that because of this multiplying effect of the weight reductions in the outer portions of the rotary elements, rotor system weight reductions of up to about 60% can be achieved if increased strength and/or reduced density goals of the rotating elements of an engine can be achieved.